Combustor assembly for a turbine engine

ABSTRACT

A combustor assembly for a gas turbine engine is provided. The combustor assembly includes a liner at least partially defining a combustion chamber and extending between an aft end a forward end generally along an axial direction. The combustor assembly also includes an annular dome including an enclosed surface defining a slot for receipt of the forward end of the liner. A cap is positioned at the forward end of the liner and at least partially positioned within the slot defined by the enclosed surface of the annular dome. The cap includes a surface configured to contact at least one of the enclosed surface of the annular dome and the forward end of the liner. Such a configuration may form a substantially airtight seal between the forward end of the liner and the annular dome despite a relative thermal expansion between the components.

CROSS-REFERENCE TO RELATED APPLICATION

This application is a division of U.S. application Ser. No. 14/842,883,filed on Sep. 2, 2015, titled “COMBUSTOR ASSEMBLY FOR A TURBINE ENGINE”,which is hereby expressly incorporated herein by reference in itsentirety.

FIELD OF THE INVENTION

The present subject matter relates generally to a gas turbine engine, ormore particularly to a combustor assembly for a gas turbine engine.

BACKGROUND OF THE INVENTION

A gas turbine engine generally includes a fan and a core arranged inflow communication with one another. Additionally, the core of the gasturbine engine general includes, in serial flow order, a compressorsection, a combustion section, a turbine section, and an exhaustsection. In operation, air is provided from the fan to an inlet of thecompressor section where one or more axial compressors progressivelycompress the air until it reaches the combustion section. Fuel is mixedwith the compressed air and burned within the combustion section toprovide combustion gases. The combustion gases are routed from thecombustion section to the turbine section. The flow of combustion gassesthrough the turbine section drives the turbine section and is thenrouted through the exhaust section, e.g., to atmosphere.

More commonly, non-traditional high temperature materials, such asceramic matrix composite (CMC) materials, are being used as structuralcomponents within gas turbine engines. For example, given an ability forCMC materials to withstand relatively extreme temperatures, there isparticular interest in replacing components within the combustionsection of the gas turbine engine with CMC materials. More particularly,an inner liner and an outer liner of gas turbine engines are morecommonly being formed of CMC materials.

However, certain gas turbine engines have had problems accommodatingcertain mechanical properties of the CMC materials incorporated therein.For example, CMC materials have different coefficients of thermalexpansion than the traditional metal materials. Accordingly, couplingthe CMC materials to the traditional metal materials can be problematic.For example, special care must be taken in attaching the inner liner andouter liner to a metallic inner dome structure and a metallic outer domestructure, respectively.

Moreover, certain gas turbine engines having the inner and outer linersformed of CMC materials have difficulty in controlling an amount ofhigh-pressure air that flows through one or more connection points—e.g.,between the inner liner and inner dome structure and the outer liner andouter dome structure—into a combustion chamber at least partiallydefined by the inner and outer liners.

Accordingly, a combustor assembly having one more features allowing fora CMC liner to be attached to a respective metallic dome structure at anattachment point while controlling an amount of airflow therethroughwould be useful. More particularly, a combustor assembly having one morefeatures allowing for a CMC liner to be attached to a respectivemetallic dome structure at an attachment point while controlling anamount of airflow therethrough and allowing for relative thermalexpansion would be particularly beneficial.

BRIEF DESCRIPTION OF THE INVENTION

Aspects and advantages of the invention will be set forth in part in thefollowing description, or may be obvious from the description, or may belearned through practice of the invention.

In one exemplary embodiment of the present disclosure, a combustorassembly for a gas turbine engine is provided. The combustor assemblydefines an axial direction and includes a liner at least partiallydefining a combustion chamber. The liner extends between an aft end anda forward end generally along the axial direction. The combustorassembly also includes an annular dome including an enclosed surfacedefining a slot for receipt of the forward end of the liner. Thecombustor assembly also includes a cap positioned at the forward end ofthe liner and at least partially positioned within the slot defined bythe enclosed surface of the annular dome. The cap includes a surfaceconfigured to contact at least one of the enclosed surface of theannular dome or the forward end of the liner.

In another exemplary embodiment of the present disclosure, a capassembly for a liner of a gas turbine engine combustor assembly isprovided. The cap assembly includes a first arm and a second armextending substantially parallel with the first arm. The first andsecond arms together define an opening for receipt of a forward end ofthe liner. The cap assembly also includes a base extending between thefirst and second arms and defining an inside surface and an outsidesurface. The cap assembly also includes a resilient member positionedadjacent to the inside surface of the base for pressing the base awayfrom the forward end of the liner and forming a seal between the baseand the forward end of the liner when the cap assembly is positionedover the forward end of the liner.

In still another exemplary embodiment of the present disclosure, a gasturbine engine defining an axial direction is provided. The gas turbineengine includes a compressor section, a turbine section mechanicallycoupled to the compressor section through a shaft, and a combustorassembly disposed between the compressor section and the turbinesection. The combustor assembly includes a liner at least partiallydefining a combustion chamber and extending between an aft end and aforward end generally along the axial direction. The combustor assemblyalso includes an annular dome including an enclosed surface defining aslot for receipt of the forward end of the liner. The combustor assemblyalso includes a cap positioned at the forward end of the liner and atleast partially positioned within the slot defined by the enclosedsurface of the annular dome. The cap includes a surface configured tocontact at least one of the enclosed surface of the annular dome or theforward end of the liner.

These and other features, aspects and advantages of the presentinvention will become better understood with reference to the followingdescription and appended claims. The accompanying drawings, which areincorporated in and constitute a part of this specification, illustrateembodiments of the invention and, together with the description, serveto explain the principles of the invention.

BRIEF DESCRIPTION OF THE DRAWINGS

A full and enabling disclosure of the present invention, including thebest mode thereof, directed to one of ordinary skill in the art, is setforth in the specification, which makes reference to the appendedfigures, in which:

FIG. 1 is a schematic cross-sectional view of an exemplary gas turbineengine according to various embodiments of the present subject matter.

FIG. 2 is a perspective, cross-sectional view of a combustor assembly inaccordance with an exemplary embodiment of the present disclosure.

FIG. 3 is a schematic, cross-sectional view of the exemplary combustorassembly of FIG. 2.

FIG. 4 is a close up, cross-sectional view of an attachment point of theexemplary combustor assembly of FIG. 2, where a forward end of an outerliner is attached to an outer annular dome.

FIG. 5 is a close-up, cross-sectional view of an attachment point of acombustor assembly in accordance with another exemplary embodiment ofthe present disclosure, where a forward end of an outer liner isattached to an outer annular dome.

DETAILED DESCRIPTION OF THE INVENTION

Reference will now be made in detail to present embodiments of theinvention, one or more examples of which are illustrated in theaccompanying drawings. The detailed description uses numerical andletter designations to refer to features in the drawings. Like orsimilar designations in the drawings and description have been used torefer to like or similar parts of the invention. As used herein, theterms “first”, “second”, and “third” may be used interchangeably todistinguish one component from another and are not intended to signifylocation or importance of the individual components. The terms“upstream” and “downstream” refer to the relative direction with respectto fluid flow in a fluid pathway. For example, “upstream” refers to thedirection from which the fluid flows, and “downstream” refers to thedirection to which the fluid flows.

Referring now to the drawings, wherein identical numerals indicate thesame elements throughout the figures, FIG. 1 is a schematiccross-sectional view of a gas turbine engine in accordance with anexemplary embodiment of the present disclosure. More particularly, forthe embodiment of FIG. 1, the gas turbine engine is a high-bypassturbofan jet engine 10, referred to herein as “turbofan engine 10.” Asshown in FIG. 1, the turbofan engine 10 defines an axial direction A(extending parallel to a longitudinal centerline 12 provided forreference) and a radial direction R. In general, the turbofan 10includes a fan section 14 and a core turbine engine 16 disposeddownstream from the fan section 14.

The exemplary core turbine engine 16 depicted generally includes asubstantially tubular outer casing 18 that defines an annular inlet 20.The outer casing 18 encases, in serial flow relationship, a compressorsection including a booster or low pressure (LP) compressor 22 and ahigh pressure (HP) compressor 24; a combustion section 26; a turbinesection including a high pressure (HP) turbine 28 and a low pressure(LP) turbine 30; and a jet exhaust nozzle section 32. A high pressure(HP) shaft or spool 34 drivingly connects the HP turbine 28 to the HPcompressor 24. A low pressure (LP) shaft or spool 36 drivingly connectsthe LP turbine 30 to the LP compressor 22.

For the embodiment depicted, the fan section 14 includes a variablepitch fan 38 having a plurality of fan blades 40 coupled to a disk 42 ina spaced apart manner. As depicted, the fan blades 40 extend outwardlyfrom disk 42 generally along the radial direction R. Each fan blade 40is rotatable relative to the disk 42 about a pitch axis P by virtue ofthe fan blades 40 being operatively coupled to a suitable actuationmember 44 configured to collectively vary the pitch of the fan blades 40in unison. The fan blades 40, disk 42, and actuation member 44 aretogether rotatable about the longitudinal axis 12 by LP shaft 36 acrossa power gear box 46. The power gear box 46 includes a plurality of gearsfor stepping down the rotational speed of the LP shaft 36 to a moreefficient rotational fan speed.

Referring still to the exemplary embodiment of FIG. 1, the disk 42 iscovered by rotatable front nacelle 48 aerodynamically contoured topromote an airflow through the plurality of fan blades 40. Additionally,the exemplary fan section 14 includes an annular fan casing or outernacelle 50 that circumferentially surrounds the fan 38 and/or at least aportion of the core turbine engine 16. It should be appreciated that thenacelle 50 may be configured to be supported relative to the coreturbine engine 16 by a plurality of circumferentially-spaced outletguide vanes 52. Moreover, a downstream section 54 of the nacelle 50 mayextend over an outer portion of the core turbine engine 16 so as todefine a bypass airflow passage 56 therebetween.

During operation of the turbofan engine 10, a volume of air 58 entersthe turbofan 10 through an associated inlet 60 of the nacelle 50 and/orfan section 14. As the volume of air 58 passes across the fan blades 40,a first portion of the air 58 as indicated by arrows 62 is directed orrouted into the bypass airflow passage 56 and a second portion of theair 58 as indicated by arrow 64 is directed or routed into the LPcompressor 22. The ratio between the first portion of air 62 and thesecond portion of air 64 is commonly known as a bypass ratio. Thepressure of the second portion of air 64 is then increased as it isrouted through the high pressure (HP) compressor 24 and into thecombustion section 26, where it is mixed with fuel and burned to providecombustion gases 66.

The combustion gases 66 are routed through the HP turbine 28 where aportion of thermal and/or kinetic energy from the combustion gases 66 isextracted via sequential stages of HP turbine stator vanes 68 that arecoupled to the outer casing 18 and HP turbine rotor blades 70 that arecoupled to the HP shaft or spool 34, thus causing the HP shaft or spool34 to rotate, thereby supporting operation of the HP compressor 24. Thecombustion gases 66 are then routed through the LP turbine 30 where asecond portion of thermal and kinetic energy is extracted from thecombustion gases 66 via sequential stages of LP turbine stator vanes 72that are coupled to the outer casing 18 and LP turbine rotor blades 74that are coupled to the LP shaft or spool 36, thus causing the LP shaftor spool 36 to rotate, thereby supporting operation of the LP compressor22 and/or rotation of the fan 38.

The combustion gases 66 are subsequently routed through the jet exhaustnozzle section 32 of the core turbine engine 16 to provide propulsivethrust. Simultaneously, the pressure of the first portion of air 62 issubstantially increased as the first portion of air 62 is routed throughthe bypass airflow passage 56 before it is exhausted from a fan nozzleexhaust section 76 of the turbofan 10, also providing propulsive thrust.The HP turbine 28, the LP turbine 30, and the jet exhaust nozzle section32 at least partially define a hot gas path 78 for routing thecombustion gases 66 through the core turbine engine 16.

Referring now to FIGS. 2 and 3, close-up cross-sectional views areprovided of the combustion section 26 of the exemplary turbofan engine10 of FIG. 1. More particularly, FIG. 2 provides a perspective,cross-sectional view of a combustor assembly 100, which may bepositioned in the combustion section 26 of the exemplary turbofan engine10 of FIG. 1, in accordance with an exemplary embodiment of the presentdisclosure, and FIG. 3 provides a side, cross-sectional view of theexemplary combustor assembly 100 of FIG. 2. Notably, FIG. 2 provides aperspective, cross-sectional view of the combustor assembly 100 havingan outer combustor casing 136 removed for clarity.

As shown, the combustor assembly 100 generally includes an inner liner102 extending between and aft end 104 and a forward end 106 generallyalong the axial direction A, as well as an outer liner 108 alsoextending between and aft end 110 and a forward end 112 generally alongthe axial direction A. The inner and outer liners 102, 108 together atleast partially define a combustion chamber 114 therebetween. The innerand outer liners 102, 108 are each attached to an annular dome. Moreparticularly, the combustor assembly 100 includes an inner annular dome116 attached to the forward end 106 of the inner liner 102 and an outerannular dome 118 attached to the forward end 112 of the outer liner 108.As will be discussed in greater detail below, the inner and outerannular domes 116, 118 each include an enclosed surface 120 defining aslot 122 for receipt of the forward ends 106, 112 of the respectiveinner and outer liners 102, 108.

The combustor assembly 100 further includes a plurality of fuel airmixers 126 (FIG. 3) spaced along a circumferential direction C withinthe outer dome 118. More particularly, the plurality of fuel air mixers126 are disposed between the outer dome 118 and the inner dome 116 alongthe radial direction R. Compressed air from the compressor section ofthe turbofan engine 10 flows into or through the fuel air mixers 126,where the compressed air is mixed with fuel and ignited to create thecombustion gases 66 within the combustion chamber 114. The inner andouter domes 116, 118 are configured to assist in providing such a flowof compressed air from the compressor section into or through the fuelair mixers 126. For example, the outer dome 118 includes an outer cowl126 at a forward end 128 and the inner dome 116 similarly includes aninner cowl 130 at a forward end 132. The outer cowl 126 and inner cowl130 may assist in directing the flow of compressed air from thecompressor section 26 into or through one or more of the fuel air mixers126.

Moreover, the inner and outer domes 116, 118 each include attachmentportions configured to assist in mounting the combustor assembly 100within the turbofan engine 10. For example, the outer dome 118 includesan attachment extension 134 configured to be mounted to an outercombustor casing 136 (FIG. 3) and the inner dome 116 includes a similarattachment extension 138 configured to attach to an annular supportmember 140 (FIG. 3) within the turbofan engine 10. In certain exemplaryembodiments, the inner dome 116 may be formed integrally as a singleannular component, and similarly, the outer dome 118 may also be formedintegrally as a single annular component. It should be appreciated,however, that in other exemplary embodiments, the inner dome 116 and/orthe outer dome 118 may alternatively be formed by one or more componentsjoined in any suitable manner. For example, with reference to the outerdome 118, in certain exemplary embodiments, the outer cowl 126 may beformed separately from the outer dome 118 and attached to the forwardend 128 of the outer dome 118 using, e.g., a welding process. Similarly,the attachment extension 134 may also be formed separately from theouter dome 118 and attached to the forward end 128 of the outer dome 118using, e.g., a welding process. Additionally, or alternatively, theinner dome 116 may have a similar configuration.

Referring still to FIGS. 2 and 3, the exemplary combustor assembly 100further includes a plurality of heat shields 142 positioned around eachfuel air mixer 124, arrange circumferentially. The heat shields 142, forthe embodiment depicted, are attached to and extend between the outerdome 118 and the inner dome 116. The heat shields 142 are configured toprotect certain components of the turbofan engine 10 from the relativelyextreme temperatures of the combustion chamber 114.

For the embodiment depicted, the inner liner 102 and outer liner 108 areeach comprised of a ceramic matrix composite (CMC) material, which is anon-metallic material having high temperature capability. Exemplary CMCmaterials utilized for such liners 102, 108 may include silicon carbide,silicon, silica or alumina matrix materials and combinations thereof.Ceramic fibers may be embedded within the matrix, such as oxidationstable reinforcing fibers including monofilaments like sapphire andsilicon carbide (e.g., Textron's SCS-6), as well as rovings and yarnincluding silicon carbide (e.g., Nippon Carbon's NICALON®, UbeIndustries' TYRANNO®, and Dow Corning's SYLRAMIC®), alumina silicates(e.g., Nextel's 440 and 480), and chopped whiskers and fibers (e.g.,Nextel's 440 and SAFFIL®), and optionally ceramic particles (e.g.,oxides of Si, Al, Zr, Y and combinations thereof) and inorganic fillers(e.g., pyrophyllite, wollastonite, mica, talc, kyanite andmontmorillonite). CMC materials may have coefficients of thermalexpansion in the range of about 1.3×10⁻⁶ in/in/° F. to about 3.5×10⁻⁶in/in/° F. in a temperature of approximately 1000-1200° F.

By contrast, the inner dome 116 and outer dome 118, including the innercowl 130 and outer cowl 126, respectively, may be formed of a metal,such as a nickel-based superalloy (having a coefficient of thermalexpansion of about 8.3-8.5×10⁻⁶ in/in/° F. in a temperature ofapproximately 1000-1200° F.) or cobalt-based superalloy (having acoefficient of thermal expansion of about 7.8-8.1×10⁻⁶ in/in/° F. in atemperature of approximately 1000-1200° F.). Thus, the inner and outerliners 102, 108 may be better able to handle the extreme temperatureenvironment presented in the combustion chamber 114. However, attachingthe inner and outer liners 102, 108 to the respective inner and outerdomes 116, 118 presents a problem due to the differing mechanicalcharacteristics of the components. Accordingly, as will be discussedbelow, a plurality of specially designed mounting assemblies 144 areutilized to attach the forward end 106 of the inner liner 102 to theinner dome 116, as well as to attach the forward end 112 of the outerliner 108 to the outer dome 118. The mounting assemblies 144 areconfigured to accommodate the relative thermal expansion between theinner and outer domes 116, 118 and the inner and outer liners 102, 108,respectively, along the radial direction R.

Referring particularly to FIG. 3, at the aft end 104 of the inner liner102 and at the aft end 110 of the outer liner 108, the combustorassembly 100 includes an inner piston ring 146 and an outer piston ring148, respectively. The inner piston ring 146 is attached to an innerpiston ring holder 150 extending from and attached to an interior casing(which for the embodiment depicted is the annular support member 140).Similarly, the outer piston ring 148 is attached to an outer piston ringholder 152 extending from and attached to an outer casing (which for theembodiment depicted includes the outer combustor casing 136 and an outerturbine casing 154). The inner piston ring holder 150 and the outerpiston ring holder 152 are configured to accommodate an expansion of theinner liner 102 and the outer liner 108 generally along the axialdirection A, as well as generally along the radial direction R.

As will be discussed in greater detail below, the above configurationmay allow for the relative thermal expansions of the inner and outerliners 102, 108, formed of a CMC material, and the inner and outer domes116, 118, formed of a metal material, while controlling an airflow ofrelatively high pressure compressed air from the compressor section 26into the relatively low pressure combustion chamber 114. Moreparticularly, such a configuration may control an airflow of relativelyhigh pressure compressed air in a high pressure plenum 156 definedbetween the outer liner 108 and the outer combustor casing 136 into therelatively low pressure combustion chamber 114, as well as an airflow ofrelatively high pressure compressed air in an inner passage 158positioned radially inward from the inner liner 102 into the relativelylow pressure combustion chamber 114.

Referring particularly to FIG. 3, and as is discussed above, thecombustion gases 66 flow from the combustion chamber 114 into andthrough the turbine section of the turbofan engine 10 where a portion ofthermal and/or kinetic energy from the combustion gases 66 is extractedvia sequential stages of turbine stator vanes and turbine rotor blades.A stage 1 turbine blade 160 is depicted schematically in FIG. 3, aft ofthe compressor assembly 100.

Referring now particularly to FIG. 4, a close up, cross-sectional viewof an attachment point where the forward end 112 of the outer liner 108is attached to the outer annular dome 118 is depicted, taken alongCircle 4-4 of FIG. 3.

As stated, to allow for a relative thermal expansion of the outer liner108 and outer dome 118, the mounting assemblies 144 are providedextending through the slots 122 defined by the enclosed surfaces 120 ofthe inner and outer annular domes 116, 118. More particularly, referringspecifically to the outer dome 118 and forward end 112 of the outerliner 108 depicted in FIG. 4, the outer dome 118 includes a base plate162 and a yolk 164. The base plate 162 and the yolk 164 each extendsubstantially parallel to one another, which for the embodiment depictedis a direction substantially parallel to the axial direction A of theturbofan engine 10. Additionally, in certain exemplary embodiments, theyolk 164 may extend circumferentially with the outer dome 118, trackingthe base plate 162. With such a configuration, the slot 122 may beconsidered an annular slot. However, in other embodiments, the yolk 164may include a plurality of circumferentially spaced tabs (see FIG. 2),each of the individual tabs of the yolk 164 defining individualsegmented portions of the slot 122 with the base plate 162.

The exemplary mounting assembly 144 depicted extends through the yolk164 of the outer dome 118, through the forward end 112 of the outerliner 108 (positioned in the slot 122 defined by the outer dome 118),and through the base plate 162 of the outer dome 118. For the embodimentdepicted, the mounting assembly 144 includes a pin 166 and a bushing168. The pin 166 includes a head 170 and a body 172, the body 172extending through the yolk 164, the forward end 112 of the outer liner108 (positioned in the slot 122), and the base plate 162. A nut 174 isattached to a distal end of the body 172 of the pin 166. In certainexemplary embodiments, the pin 166 may be configured as a bolt and thenut 174 may be rotatably engaged with the pin 166 for tightening themounting assembly 144. Alternatively, however, in other exemplaryembodiments, the pen 166 and nut 174 may have any other suitableconfiguration. For example, in other exemplary embodiments, the pin 166may include a body 172 defining a substantially smooth cylindrical shapeand the nut 174 may be configured as a clip.

Additionally, the bushing 168 is generally cylindrical in shape andpositioned around the body 172 of the pin 166 within the slot 122. Thebushing 168 is pressed between the yolk 164 and the base plate 162.Moreover, for the embodiment depicted, the mounting assembly 144includes a metal grommet 176 positioned around the bushing 168 within anopening defined in the forward end 112 of the outer liner 108. The metalgrommet 176 may reduce an amount of wear on the forward end 112 of theouter liner 108 as the outer liner 108 moves inwardly and outwardlygenerally along the radial direction R relative to the outer dome 118.More particularly, the metal grommet 176 may reduce an amount of weararound an opening 177 in the outer liner through which the mountingassembly 144 extends.

Referring still to FIG. 4, the exemplary combustor assembly 100 furtherincludes a cap 178 positioned at the forward end 112 of the outer liner108. More particularly, for the embodiment depicted, the cap 178 ispositioned over the forward end 112 of the outer liner 108 and at leastpartially within the slot 122 of the outer dome 118. The cap 178generally includes a first arm 180 and a second arm 182, the second arm182 extending substantially parallel with the first arm 180. The firstand second arms 180, 182 together define an opening 184 for receipt ofthe forward end 112 of the outer liner 108. The cap 178 also includes abase 186 extending between the first and second arms 180, 182 anddefining an inside surface 188 and an outside, or end, surface 190. Thefirst arm 180, the second arm 182, and the base 186 are all formedintegrally for the embodiment depicted from a metal material. Forexample, the first arm 180, the second arm 182, and the base 186 may beformed using a casting process, or alternatively, the opening 184 may beformed between the first and second arms 180, 182 using an extrusionprocess. Alternatively, however, the cap 178 may be formed of individualarm and base components joined using, e.g., a welding process. Moreover,in certain exemplary embodiments, the cap 178 may be a single annularcomponent, or alternatively the cap 178 may be formed of a plurality ofcomponents arrange circumferentially over the forward end 112 of theouter liner 108. Further, in still other embodiments, the cap 178 may beformed partially or completely of a suitable CMC material.

Referring still to the embodiment depicted, the first and second arms180, 182 of the cap 178 extend past the mounting assemblies 144.Accordingly, the first arm 180 and the second arm 182 may each defineone or more openings for receiving at least a portion of one or more ofthe mounting assemblies 144 mounting the forward end 112 of the outerliner 108 to the outer dome 118. For example, the first and second arms180, 182 depicted may each define one or more openings allowing themetal grommet 176, the bushing 168, and the pin 166 of each mountingassembly 144 to extend therethrough.

For the exemplary embodiment depicted, the base 186 of the cap 178 andthe forward end 112 of the outer liner 108 define a gap 192 therebetweenwith a resilient member 194 positioned therein (i.e., adjacent to theinside surface 188 of the base 186 and the forward end 112 of the outerliner 108). The purpose of the resilient member 194 is twofold. First,the resilient member 194 is configured to form a seal between the insidesurface 188 of the base 186 of the cap 178 and the forward end 112 ofthe outer liner 108. Second, the resilient member 194 is configured topress the base 186 of the cap 178 away from the forward end 112 of theliner 108 such that the end surface 190 of the cap 178 is pressedagainst the enclosed surface 120 of the outer dome 118. Accordingly,such a configuration may allow the cap 178 to form a substantiallyairtight seal between the forward end 112 of the outer liner 108 and theouter dome 118.

In certain exemplary embodiments, the resilient member 194 may be a ropeseal, such as a braided rope seal having a silicone core. Alternatively,however, in other exemplary embodiments, any other suitable resilientmember 194 may be provided for pressing the base 186 of the cap 178 awayfrom the forward end of the liner and forming a seal between the insidesurface 188 of the base 186 of the cap 178 and the forward end 112 ofthe liner 108. For example, in other exemplary embodiments, theresilient member 194 may be a W-seal, a wire seal, or any other suitableseal.

Moreover, referring back to FIG. 3, it should be appreciated that theforward end 106 of the inner liner 102 may be attached to the inner dome116 in substantially the same manner that the forward end 112 of theouter liner 108 is attached to the outer dome 118. For example, a cap(similar to the cap 178 positioned over the forward end 112 of the outerliner 108) may be positioned over the forward end 106 of the inner liner102 and at least partially within the slot 122 of the inner dome 116. Anend surface of such cap may contact the enclosed surface 120 of theinner dome 116 such that a substantially airtight seal is formed betweenthe end surface of such cap and the enclosed surface 120 of the innerdome 116. Further, such a cap may define a gap (similar to the gap 192defined between the inner surface 188 of the cap 178 and the forward end112 of the outer liner 108) between an inner surface and the forward end106 of the inner liner 102 with a suitable resilient member 194positioned therein.

A combustor in accordance with an exemplary embodiment of the presentdisclosure assembly having a cap positioned over an inner liner or anouter liner may be capable of controlling an airflow from a relativelyhigh pressure plenum or a relatively high pressure inner passage into acombustion chamber through an attachment point between the inner orouter liners and an inner or outer dome. Moreover, such a combustorassembly may be capable of controlling an airflow from a relatively highpressure plenum or a relatively high pressure inner passage into acombustion chamber through an attachment point between the inner orouter liners and an inner or outer dome while still accommodating arelative thermal expansion between the inner or outer liners and inneror outer domes.

Reference will now be made to FIG. 5. FIG. 5 provides a close-up,cross-sectional view of a combustor assembly 100 in accordance withanother exemplary embodiment of the present disclosure. Moreparticularly, FIG. 5 provides a close-up, cross-sectional view of anattachment point where a forward end 112 of an outer liner 108 isattached to an outer annular dome 118. The exemplary combustor assembly100 of FIG. 5 may be configured in substantially the same manner as theexemplary combustor assembly 100 described above with reference to FIGS.2 through 4. Accordingly, the same or similar numbering refers to thesame or similar components.

As is depicted, the forward end 112 of the outer liner 108 is positionedwithin a slot 122 defined by an enclosed surface 120 of the outerannular dome 118. A mounting assembly 144 attaches the forward end 112of the outer liner 108 to the outer annular dome 118. Additionally, theexemplary combustor assembly 100 depicted in FIG. 5 includes a cap 178′positioned at the forward end 112 of the outer liner 108 and at leastpartially within the slot 122 defined by the enclosed surface 120 of theannular dome 118. The exemplary cap 178′ depicted defines an insidesurface 188′ configured to contact the forward end 112 of the liner 108and an end surface 190′ positioned opposite the inside surface 188′. Theend surface 190′ defines a notch 196. A resilient member 194 ispositioned adjacent to the end surface 190′ of the cap 178′, within thenotch 196. The cap 178′ and resilient member 194 are configured to forma seal between the end surface 190′ of the cap 178′ and the enclosedsurface 120 of the annular dome 118, as well as between the insidesurface 188′ of the cap 178′ and the forward end 112 of the liner 108.More particularly, the resilient member 194 is configured to form a sealbetween the enclosed surface 120 of the annular dome 118 and the endsurface 190′ of the cap 178′, and is also configured to press the cap178′ against the forward end 112 of the outer liner 108, such that theinside surface 188′ of the cap 178′ contacts the forward end 112 of theouter liner 108. Such a configuration may allow for the exemplarycombustor assembly 100 depicted to control an airflow through theattachment point between the outer annular dome 118 and the outer liner108 as the outer annular dome 118 thermally expands relative to theouter liner 108 along the radial direction R.

This written description uses examples to disclose the invention,including the best mode, and also to enable any person skilled in theart to practice the invention, including making and using any devices orsystems and performing any incorporated methods. The patentable scope ofthe invention is defined by the claims, and may include other examplesthat occur to those skilled in the art. Such other examples are intendedto be within the scope of the claims if they include structural elementsthat do not differ from the literal language of the claims, or if theyinclude equivalent structural elements with insubstantial differencesfrom the literal languages of the claims.

1. A combustor assembly for a gas turbine engine defining an axialdirection, the combustor assembly comprising: a liner at least partiallydefining a combustion chamber and extending between an aft end and aforward end generally along the axial direction; an annular domeincluding an enclosed surface defining a slot for receipt of the forwardend of the liner; and a cap positioned at the forward end of the linerand at least partially positioned within the slot defined by theenclosed surface of the annular dome, the cap including a surfaceconfigured to contact at least one of the enclosed surface of theannular dome or the forward end of the liner; wherein the surface of thecap is an inside surface configured to contact the forward end of theliner, wherein the cap further defines an end surface opposite theinside surface, and wherein a resilient member is positioned adjacentthe end surface of the cap for forming a seal between the end surface ofthe cap and the enclosed surface of the annular dome.
 2. A cap assemblyfor a liner of a gas turbine engine combustor assembly, the cap assemblycomprising: a first arm; a second arm extending substantially parallelwith the first arm, the first and second arms together defining anopening for receipt of a forward end of the liner; a base extendingbetween the first and second arms and defining an inside surface and anoutside surface; and a resilient member positioned adjacent to theinside surface of the base for pressing the base away from the forwardend of the liner and forming a seal between the base and the forward endof the liner when the cap assembly is positioned over the forward end ofthe liner.
 3. The cap assembly of claim 2, wherein the first arm, thesecond arm, and the base are all formed integrally of a metal materialor a CMC material.
 4. The cap assembly of claim 2, wherein the resilientmember is a rope seal including a silicone core.
 5. The cap assembly ofclaim 2, wherein the first arm and the second arm each define one ormore openings for receiving a mounting assembly for mounting a forwardend of the liner to an annular dome.
 6. A gas turbine engine defining anaxial direction, the gas turbine engine comprising: a compressorsection; a turbine section mechanically coupled to the compressorsection through a shaft; and a combustor assembly disposed between thecompressor section and the turbine section, the combustor assemblyincluding a liner at least partially defining a combustion chamber andextending between an aft end and a forward end generally along the axialdirection; an annular dome including an enclosed surface defining a slotfor receipt of the forward end of the liner; and a cap positioned at theforward end of the liner and at least partially positioned within theslot defined by the enclosed surface of the annular dome, the capincluding a surface configured to contact at least one of the enclosedsurface of the annular dome or the forward end of the liner.
 7. The gasturbine engine of claim 6, wherein the cap is positioned over theforward end of the liner, and wherein the surface of the cap is an endsurface configured to contact the enclosed surface of the annular dome8. The gas turbine engine of claim 7, wherein a resilient member ispositioned in a gap defined between the cap and the forward end of theliner, the resilient member configured to press the end surface of thecap against the enclosed surface and form a seal between the cap and theliner.
 9. The gas turbine engine of claim 6, wherein the gas turbineengine further defines a circumferential direction, and wherein thecombustor assembly further includes a plurality fuel/air mixers spacedalong the circumferential direction within the annular dome.
 10. Thecombustor assembly of claim 1, wherein the cap is positioned over theforward end of the liner, and wherein the first surface of the cap is anend surface configured to contact the enclosed surface of the annulardome.
 11. The combustor assembly of claim 1, wherein the end surfacedefines a notch.
 12. The combustor assembly of claim 11, wherein theresilient member is positioned within the notch.
 13. The combustorassembly of claim 1, wherein the resilient member is configured to pressthe cap against the forward end of the liner such that the insidesurface of the cap contacts the forward end of the liner.
 14. Thecombustor assembly of claim 1, wherein the resilient member is a ropeseal including a silicone core.
 15. The combustor assembly of claim 1,wherein the liner is an outer liner and wherein the annular dome is anouter annular dome.
 16. The combustor assembly of claim 1, wherein theliner is an inner liner and wherein the annular dome is an inner annulardome.
 17. The combustor assembly of claim 1, wherein the liner iscomprised of a ceramic matrix composite material.
 18. The combustorassembly of claim 17, wherein the annular dome is comprised of a metalmaterial.
 19. The combustor assembly of claim 18, wherein the annulardome includes a base plate and a yoke, the base plate and the yokeextending substantially parallel to one another, the enclosed surface ofthe annular dome including a second surface of the base plate and athird surface of the yoke.
 20. The combustor assembly of claim 19,further comprising: a mounting assembly extending through the yoke ofthe annular dome, through the forward end of the liner, and through thebase plate of the annular dome, wherein the mounting assembly attachesthe liner to the annular dome.